This invention relates to centrifugal impellers and, in particular, concerns centrifugal impellers used as final compressor stages in aircraft gas turbine engines.
Gas turbine engines have become the standard propulsion source for all types of aircraft but the very smallest. The aircraft gas turbine engine operates as an open Brayton cycle system whose work output is either in the form of a high energy exhaust or a mechanical shaft rotation. In gas turbine engines that produce a high energy exhaust stream, turbofan or turbojet engines, the compressors are generally of the axial type. Axial compressors are made up of a series of rotating stages of axial blades which compress inlet air prior to combustion. In gas turbine engines of the rotating shaft type, turboshaft or turboprop engines, axial-centrifugal compressors have been found to be more suited in many applications. The axial-centrifugal compressors have a series of axial stages similar to pure axial compressors prior to a final centrifugal stage situated just upstream of a gas diffuser and a combustor. The centrifugal stage provides a high level of compression in a minimum of length.
The centrifugal impellers used in the axial-centrifugal compressors have been found lacking in several respects in comparison with axial stages.
In recent years the centrifugal impellers devised have not been able to achieve the efficiency of axial stages. This shortcoming in efficiency has limited centrifugal impeller use in turbojet engines and hindered axial-centrifugal compressor efficiency in turboshaft type engines.